The invention relates to the field of guidance, navigation, and control avionics. In particular the invention relates to inertial attitude sensors.
Attitude determination is crucial to the flight of both aeronautical and space vehicles. In particular, attitude determination for both spinning and three-axis stabilized satellites is a critical operational function.
However, a choice must typically be made between the accuracy of the attitude determination and the size, weight, power and computational requirements of the associated system or methods. Although attitude determination is critical, these are not simple choices because space, weight, power and computational resources are at a premium in aeronautical and space vehicles. Moreover, space, weight, power and computational resources become even more scarce as the size of the aeronautical or space vehicle decreases.
A need therefore exists for apparatus and methods that can provide sufficiently accurate attitude determination with reduced space, weight, power and/or computational requirements.
In one aspect, the invention provides methods for measuring the inertial attitude of an aeronautical or space vehicle (hereinafter collectively referred to as xe2x80x9caerospace vehiclexe2x80x9d). In one embodiment, the methods provide the inertial attitude for a spinning aerospace vehicle. In another embodiment, the methods provide the inertial attitude of a substantially non-spinning (e.g., 3-axis stabilized) aerospace vehicle.
The methods employ a rotational astronomical object-sighting concept wherein radii measurements of curved astronomical object tracks of three or more astronomical objects observed by an optical sensor in the coordinate reference frame of an aerospace vehicle, are used to obtain the direction of an axis of the vehicle. In one embodiment, the methods determine the direction, or inertial attitude, of an axis in the right ascension/declination reference frame. In another embodiment, the methods of the invention further determine the roll angle of the aerospace vehicle with respect to an axis, in addition to the inertial attitude of the axis. In general, the invention eliminates the need to implement complex drag-back schemes that use inertial sensor (e.g. gyroscopes) measured motion to drag-back the image of the astronomical object over the astronomical object sighting scan. In addition, the methods of the invention eliminate the need to identify the individual astronomical objects in order to determine the inertial attitude of an axis.
In other embodiments, the methods of the invention use inertial sensor information to enhance the accuracy of the determination of the inertial attitude of an axis by using the inertial sensor information to compensate for nutation of the aerospace vehicle.
In various embodiments, the methods are adapted for a spinning aerospace vehicle. In one embodiment, the aerospace vehicle spins at a rate of 20 rpm However, it is to be understood that the spin rate may be substantially higher or lower than 20 rpm, without deviating from the scope of the invention. In the embodiments adapted for a spinning aerospace vehicle, the field-of-view (FOV) of the optical sensor is oriented along the spin vector, which preferably is directed towards a substantially fixed location in the celestial sphere such as, for example, the ecliptic pole of the solar system. The star field near the spin axis may be known a priorixe2x80x94based on the aerospace vehicle missionxe2x80x94thus simplifying any star pattern recognition task or it may be unknown, the so-called xe2x80x9clost-in-spacexe2x80x9d scenario. The detected astronomical objects track out a circular arc in the field-of-view about the spin axis, which is substantially centered in the field-of-view. The signal associated with the astronomical object track, or tracks, may be integrated over time to improve the signal-to-noise ratio. In this configuration, the spin axis of the optical sensor is relatively stable in inertial space so that putative pitch/yaw motion measured by an inertial sensor (e.g., such as a MEMS gyroscope) must arise substantially from inertial sensor drift. A processor, using the measured putative pitch/yaw motion, can measure this drift and compensate for it. In another embodiment, the inertial sensor (gyroscope) drift is also updated by an external measurement, such as a sun sensor or earth horizon sensor, to provide a roll angle update. However, as noted previously, roll angle update is not critical to or for the radii measurements.
In other embodiments, the method is adapted for a non-spinning aerospace vehicle (e.g., a 3-axis stabilized satellite). In various versions of these embodiments the curved astronomical object tracks are generated by servo-controlled optical sensor rotational motion rather than a free spinning aerospace vehicle.
The invention provides several advantages to the field of inertial attitude sensing and determination. For example, the rotational astronomical object-sighting concept of the invention eliminates the need for drag-back yet provides a substantially equivalent integration of astronomical object sighting data compared to that traditionally obtained with drag-back procedures. Eliminating drag-back simplifies inertial attitude determination and reduces the need for on-board processing and/or the need for data downlinking. The rotational astronomical object-sighting concept also reduces the requirement for very accurate knowledge of the roll gyroscope (about the spin axis) scale factor error. The radii measurements are substantially orthogonal to the roll gyroscope error and therefore are substantially unaffected by the roll gyro scale factor error accumulation during large angle rotations. As a result, the methods of the invention increase the viability of using less-accurate gyroscopes (which are typically smaller, lighter and require less power than high accuracy gyroscopes) in an attitude senor without a significant performance penalty. Accordingly, various embodiments of the invention provide an inertial attitude sensor, or sensor suite, that combines small size, low power requirements and simplicity of operation with high accuracy inertial attitude determination.
In another aspect, the invention provides an attitude determination system to determine the inertial attitude of an axis of an aerospace vehicle. In one embodiment, the system comprises an optical sensor, a first memory element to store the radii of astronomical object tracks, a second memory element to store the coordinates of a plurality of astronomical objects, and an attitude processor adapted to determine the attitude of an axis.
In another aspect, the invention provides an attitude measurement apparatus comprising a high sensitivity optical sensor and a low power inertial sensor. Although low power inertial sensors typically have only modest performance (e.g., drift, scale factor error, alignment, etc), the invention provides a system synergy where the modest performance of the inertial sensor is enhanced by the high sensitivity optical sensor, which is adapted to operate as a star tracker (stellar camera), and the inertial sensor is adapted to increase the accuracy of the astronomical object tracking operation of the optical sensor. In one embodiment, the optical sensor comprises an electron bombarded charge coupled device (EBCCD) and the inertial sensor comprises a microelectromechanical system (MEMS) gyroscope. In another embodiment, the invention further comprises calibration algorithms and/or Kalman filters that can model the temperature disturbance of the inertial sensor and further improve inertial sensor performance.
In one embodiment, the high sensitivity optical sensor can track dim astronomical objects (e.g., those of relative magnitude 8 or higher) and is not limited to tracking brighter objects. Accordingly, the invention can use a small field-of-view (FOV) for the optical sensor. Limiting the FOV decreases the number of astronomical objects candidates and simplifies astronomical object pattern recognition. Simplifying astronomical object recognition reduces data processing requirements, which in turn reduces power requirements. This further simplifies astronomical object pattern recognition and also mitigates inertial sensor errors such as gyroscope scale factor.
In another embodiment, a large optical sensor FOV is used to accommodate the multiple astronomical objects (e.g., astronomical objects with relatively bright stellar magnitudes). Using a large FOV means that larger system errors can be corrected for using the astronomical object sighting. Such errors include, but are not limited to, aerospace vehicle tilt, gyroscope drift, optical sensor and/or gyroscope misalignment, launch navigation errors, and launch azimuth errors.
In another aspect, the invention provides an article of manufacture where the functionality of one or more of the methods of the invention are embedded on a computer-readable program means, such as, but not limited to, a floppy disk, a hard disk, an optical disk, a magnetic tape, a PROM, an EPROM, CD-ROM, or DVD-ROM.